1. Field of the Invention
The present invention relates to load sensing systems and more particularly to a load sensing system particularly adaptable for use with aircraft actuation systems for providing load path integrity monitoring.
2. Description of the Related Art
Modern aircraft have horizontal stabilizers located at the rear tail section of the fuselage or the forward section that are pivotally supported relative to the airplane fuselage to “trim” the aircraft during flight by selective adjustment by the operator or auto-pilot from an internal control unit. The stabilizer actuator is a variable length structural link connecting the horizontal stabilizer control surface to the fuselage structure and used to control the pitch (attitude) of the aircraft during take off, cruise and landing phases under different aerodynamic loading conditions. The stabilizer actuator is also used to recover the aircraft during severe aircraft stall situations. In this regard the stabilizer is traditionally pivotally connected to the rear section or tail section of the fuselage.
One common trimmable horizontal stabilizer actuator consists of a primary ball nut assembly connected with an actuating drive gimbal which is pivotally connected to one end of the horizontal stabilizer structure. The ball nut assembly includes a ball nut housing and a rotatable ball screw extending axially and usually vertically through the ball nut housing and a drive gimbal housing. The ball nut housing is connected to the drive gimbal housing by a trunnion segment. The ball screw, in turn, has its upper end remote from the actuating drive gimbal and is fixed from translation or axial movement by a connection to a second, support gimbal which is pivotally secured to the vertical stabilizer section or the tail section. As the ball screw is rotated, the drive gimbal will be moved in translation relative to it. Thus as the ball screw is rotated in one direction, the leading edge of the horizontal stabilizer is pivoted upward, whereas by rotating the ball screw in the other direction, the leading edge of the horizontal stabilizer is pivoted downward. Rotation of the ball screw is routinely done by a motor (electric or hydraulic, depending on system architecture) and associated gearing which is connected to the second, fixed support gimbal and which is actuated by the operator or pilot by the internal control unit. The connection of the stabilizer actuator to the stabilizer is located within the vertical stabilizer or fuselage tail section and not directly in the air stream.
The horizontal stabilizer movement, as controlled by the operator or auto-pilot, is transmitted by the ball screw through the actuating drive gimbal by way of the primary ball nut assembly which defines a primary load path. The movement has a load with tensile and compressive components as well as a torque component due to the ball screw thread lead. Failures of the primary load path such as caused by the shearing off of the connecting trunnion segment, ball screw disconnect or by the loss of nut ball members from the ball nut assembly can result in the complete loss of control of the horizontal stabilizer. However, stabilizer actuators have frequently been provided with a secondary load path for alternate control of the stabilizer and structural integrity, as well as to meet the required level of safety (failure of single load path actuator has a catastrophic outcome on the aircraft). In such structures, the primary load path is normally controllably actuated by the operator and is thus under load while the secondary load path is normally unactuated and thus unloaded. In the event of a primary load path failure, the secondary load path is automatically mobilized whereby the stabilizer actuator can continue to be controllably actuated by the operator or pilot by the internal control unit to control the position of the stabilizer. The transition of control to the secondary load path can occur quite rapidly whereby failure of the primary load path is not necessarily detected by the operator or pilot.
However, in the event of a subsequent failure of the secondary load path through continued, periodic use, control of the stabilizer will be completely lost which could result in erratic, oscillatory movement of the stabilizer whereby the ability of the pilot or operator to control the aircraft could be substantially inhibited. This problem is addressed by the present invention.
This problem was addressed in, for example, U.S. Pat. No. 6,672,540, entitled “Actuator For Aircraft Stabilizers With a Failure Responsive Lock Control Mechanism,” issued to M. A. Shaheen et al, and assigned to the present assignee, that discloses a horizontal stabilizer actuator for a winged aircraft which is selectively pivotally controlled by a pilot or operator at a remote location in the aircraft and which has a primary load path responsive to the selective control by the pilot or operator for setting the pivoted position of the stabilizer and which has a secondary load path which is responsive to a failure in the primary load path to be automatically actuated to a condition locking the stabilizer in a fixed position. The solution offered by the '540 patent offers detection of the actuator secondary load path engagement by means of a lock that will trigger and jam the ball screw in the event of a primary load path failure and secondary lock engagement. The jammed and immobilized actuator will stall the drive motor which is detected and annunciated by the system controller, thus annunciating the fault in the actuator. This solution is feasible when dealing with jamming devices needed to generate lock torque less than 3000 in-lb, which is feasible for actuators used on regional jets and small corporate jets. This solution becomes limited when dealing with large body aircraft. With large body aircraft the torque needed to stall the drive motor by means of locking and jamming the ball screw is extremely high (more than 10,000 in-lb) and is infeasible and unachievable without hugely impacting the envelope (volume) and weight of the actuator (especially the ball screw gimbal area where the lock is mounted).
The trend in aviation regulations shows that it is ever more desirable to have intelligent systems be aware in real time of the state of certain active elements, such as control surfaces. U.S. Pat. No. 7,299,702, entitled “Apparatus For Monitoring An Aircraft Flap and Application of a Dynamometric Rod,” issued to F. Gibert, discloses and claims using a dynamometric pin for replacing a flap hinge pivot. The pin differs from dynamometric pins or shafts known in the previous prior art firstly because its sensing elements are enclosed inside the body of the pin, which is itself closed by packing, so they are not subjected to bad weather, and secondly because they possess the external characteristics of the pivots presently in use for mounting flap arms, and so they require no modification to the flaps, to the arms, or to the wing. However, the '702 device is deficient because it assumes that the surface is directly mounted to the aircraft structure and does not address the need to monitor the integrity of the load paths of actuators that actuate the control surfaces and that are typically mounted between the aircraft structure and the respective control surfaces. This solution maybe applicable to very light aircraft where the control surface is directly connected to the aircraft structure and it is manually controlled by cables, pulleys and control rod without actuators.
The dynamometric pin disclosed in the '702 patent is substantially that disclosed in U.S. Pat. No. 3,695,096, entitled “Strain Detecting Load Cell,” issued to A. U. Kutsay. The '096 patent discloses and claims a strain detecting load cell which is adapted to replace or be readily interchangeable with coupling member such as a pin or bolt. The working dimensions of the cell and the part replaced are the same, except that the cell has short zones of slightly decreased diameter so that the shear strains are concentrated in these zones. An axial bore in the cell contains electrical strain gauges attached to its circumferential wall within the concentrating zones and having leads for connection to exterior measuring instrumentation such as Wheatstone bridge equipment. The arrangement and orientation of the gauges in the concentrating zones permit evaluation of the applied load both as to magnitude and direction.
U.S. Ser. No. 12/150,365, filed on Apr. 28, 2008, entitled “Actuator Load Path Monitoring System”, by Abbas M. Charafeddine et al, and assigned to present assignee discloses an actuator load path monitoring system for an aircraft having an aircraft structure, a control surface, and, an actuator connected between the aircraft structure and the control surface to support and position the control surface as desired relative to the aircraft structure. The actuator is of a type including a) an upper actuator assembly securely connected to the aircraft structure, including a motor assembly and gear assembly; b) a ball screw assembly operatively connected to the gear assembly; c) a tie-rod assembly (also called a safety-rod or safety-bar assembly) positioned within the ball screw assembly; and, d) a lower actuator assembly securely connected to the control surface, wherein actuation of the ball screw provides selected positioning of the control surface. The actuator load path monitoring system includes an upper load sensing assembly positioned in an upper load path between the upper actuator assembly and the aircraft structure. The upper load sensing assembly provides upper indications of the applied forces in the upper load path when the upper load path is disconnected. A lower load sensing assembly is positioned in a lower load path between the lower actuator assembly and the control surface. The lower load sensing assembly provides upper indications of the applied forces in the lower load path when the lower load path is disconnected. A computer system receives the upper and lower indications of applied forces and analyzes the upper and lower indications, thereby monitoring the structural integrity and safety of the upper and lower load paths by annunciating the detection of a failed portion thereof.
The Charafeddine et al patent application also discloses use of a dynamometric pin such as disclosed in Gibert and Kutsay. In all of the above cases the load sensing assemblies are not provided in an integrated fashion. In other words, the power supply conversion, logic, and enunciation modules are not incorporated into a single housing with the dynamometric pin. Also, for aircraft applications there is a need for in situ testing of the load sensing assembly because of the catastrophic consequences of sensor dormancy. As will be disclosed below, the systems and methods described herein solve these aforementioned problems.